For this type of application, there are three families of technologies, depending on the physical state of the propellants used: solid propulsion, wherein the propellant is stored in a combustion chamber, liquid propulsion, which can use one, two or even more propellants, wherein propellants must be transferred from storage tanks to a combustion chamber, and hybrid propulsion, which uses a liquid propellant and a solid propellant, and wherein a liquid propellant must be transferred to a combustion chamber in which a solid propellant is stored.
The present invention relates more precisely to devices for transferring liquid propellants to the combustion chamber and more specifically to the drive system for this transfer.
In order to be capable of providing high thrust, rocket engines must run at a high pressure of several tens of bar, for example from 30 to 50 bar for Ariane engines, with a high flow of matter.
In the case of liquid propulsion, it is the propellant feed system that must provide this flow and this pressure. Two methods are commonly used to produce this pressurized feed: direct pressurization of the propellant tanks and pumping with pumps from a low-pressure tank.
The first solution has the virtue of simplicity, but requires tanks capable of withstanding high pressures, which creates problems in terms of mass and safety. This solution is limited in practice to low-power engines, such as attitude control engines or the upper stages of launchers for example, where installing an external means of pressurization is less advantageous.
The second solution requires the use of specific pumps capable of producing the substantial flow required by the engines. This flow, combined with the huge increase in pressure required, results in pumps of considerable power, from several hundred kilowatts to several megawatts.
In current and past space launchers, these pumps are systematically driven by centrifugal turbines, generally using the same propellants as the main engine.
These centrifugal turbines are driven by hot gases. These hot gases are generally produced by taking a portion of the propellants for the rocket engine and burning these portions in a specific small combustion chamber. They can also be produced by a gas generator, often a small powder rocket.
A centrifugal turbine/pump assembly is called a turbopump. A turbopump is a complex, fragile object because it must transmit very high levels of power—several megawatts—using very high rotation speeds, for example from 10,000 to 30,000 rpm, which exerts very high mechanical stresses on the materials.
In addition, being driven by hot gases resulting from combustion produces very high temperatures in the turbine and very large temperature gradients in the transmission shafts between the turbine and the pump.
This thermal gradient effect is further accentuated when the propellants are cryogenic, the pump temperature being several tens of degrees Kelvin while only a few centimeters away, the temperature of the driving centrifugal turbine is more than 1,000 degrees Celsius.
Lastly, because of these extreme operating conditions, starting a turbopump is difficult, with one part being cooled, the other being heated, and the assembly being brought to rotation gradually enough not to cause an even higher transient gradient capable of rupturing the turbopump.
Ultimately, a turbopump is a very expensive object with a short life, used in conventional launchers which have a short operating life that is measured in minutes.
In reusable launchers like the space shuttle, the turbopumps must be replaced for nearly every flight, which is quite onerous in terms of maintenance costs.
One proposed solution for replacing a turbopump is described in the document U.S. Pat. No. 6,457,306.
This document specifically describes replacing the drive turbine of the pump with a battery-powered electric motor.
Thus, there is no longer a need for a small rocket engine driving a turbine, less propellant is consumed, there are no longer such high temperature gradients, and the assembly is more reliable and better adapted to a reusable launcher.
It is also possible to adjust the rotation of the electric motor and thus vary the propellant flows, and hence the thrust, more easily; it is also easier to control the start of the pump so as to prevent excessively high transient gradients.
On the other hand, the energy source that powers the engine must be capable of supplying a power that is measured in megawatts during the thrust phase, which entails significant mass and size constraints for this energy source and for the means for powering the electric motor.
The energy storage system and motor are ultimately very heavy.